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Topic: Basic Rocket Science Q & A (Read 307363 times)

Atlas (Mercury) and Titan (Gemini) were last gasp measures to catch up with the Soviet program in a *big hurry*.

Atlas needed thicker tankage/skin and used a "belly band" on John Glenn's first orbital Mercury. This flight also had engines that were not baffled and prone to 1 in 5 combustion instability (i.e.LOM or LOC).

The Titan had an extremely steep trajectory with 7g load on the crew and NO abort option. Steep trajectory was from underpowered booster for the given task so you can't just put an abort tower in the mix there.

John Young mentions the Gemini/Titan ascent in the special features of the In The Shadow of the Moon DVD, by saying, "I don't think comfort is what your looking for going uphill".

Neither one of these was a good manned launcher (much less optimal) so the comparison is not that valid.

These boosters were *only* "manrated" for the extremes of the Cold War/Space Race.

Each one of these crash programs were never intended to last more than a couple of years. Recall the "gap" from the end of Gemini in November of 1966 until Apollo 7 in October of 1968.

It has to do with idealised models of the expansion process in the nozzle. Shifting equilibrium assumes chemical reactions continue to take place during the expansion process, while frozen assumes that mole fractions remain constant during the expansion process. If the reaction is very fast then Isp may be closer to shifting equilibrium Isp, otherwise it will be closer to the frozen value.

It has to do with idealised models of the expansion process in the nozzle. Shifting equilibrium assumes chemical reactions continue to take place during the expansion process, while frozen assumes that mole fractions remain constant during the expansion process. If the reaction is very fast then Isp may be closer to shifting equilibrium Isp, otherwise it will be closer to the frozen value.

Adding to that, frozen underpredicts. For BOE estimates, I will use frozen. Shifting will always overpredict to an extent, because you are getting back energy from dissociated species recombining (H monatomic, rejoins to become H2, etc).

Also, if you want to try playing with the chemistry, CEA is an excellent resource. They have an online version there, but it is kind of buggy. The download is here. It's not a tool that can be used completely blind, but it is pretty user friendly, as these things go.

Not sure if this is the correct location for this question, but can anyone point me towards information about the thermal environment on the side of a cylindrical rocket during ascent?

I know that the cone of the payload fairing will get heated and I understand that the base of the vehicle will. I also note that the plume can climb back up the side of the vehicle due to low pressures in that region, but I'm really just trying to get a feel for the heating involved above that, say on the cylindrical body of the payload fairing or the upper stage tanking. Does the airflow in that area cause noticeable heating?

On a SpaceX thread the possibility of using gaseous hydrogen dissolved in liquid methane came up. It looks as if this is impractical since the solubility is far too low. Would it be possible to use slush methane in liquid hydrogen and if so would this be more practical than slush hydrogen?

On a SpaceX thread the possibility of using gaseous hydrogen dissolved in liquid methane came up. It looks as if this is impractical since the solubility is far too low. Would it be possible to use slush methane in liquid hydrogen and if so would this be more practical than slush hydrogen?

Could be possible, a good way to increase density.I see a problem pumping the liquid/solid mixture (methane is rock solid at LH temperatures).Some way to avoid damage to impellers of pumps must be studied.

On a SpaceX thread the possibility of using gaseous hydrogen dissolved in liquid methane came up. It looks as if this is impractical since the solubility is far too low. Would it be possible to use slush methane in liquid hydrogen and if so would this be more practical than slush hydrogen?

I'm struggling a bit to understand the advantage of doing this. Is the hope that dissolved methane would add little to the volume, thereby increasing volume-specific impulse (at the cost of weight-specific impulse)?

I'm struggling a bit to understand the advantage of doing this. Is the hope that dissolved methane would add little to the volume, thereby increasing volume-specific impulse (at the cost of weight-specific impulse)?

I believe the desire is for a higher, and tailorable density specific impulse.

On a SpaceX thread the possibility of using gaseous hydrogen dissolved in liquid methane came up. It looks as if this is impractical since the solubility is far too low. Would it be possible to use slush methane in liquid hydrogen and if so would this be more practical than slush hydrogen?

I'm struggling a bit to understand the advantage of doing this. Is the hope that dissolved methane would add little to the volume, thereby increasing volume-specific impulse (at the cost of weight-specific impulse)?

Coming from Elon's Reddit AMA video responses:"We can certainly improve on the chemical propulsion that has been done thus far, and I think probably a very high efficiency light hydrocarbon that uses predominantly methane is probably a good way to go, and I think that's something SpaceX will end up working on."

I was wondering what Elon was meaning and my idea was:Obtaining a (still unknown) mixture of liquid hydrocarbons that enables storing hydrogen at mild cryo temperature and low pressure would be a great advantage.Obviously adding solid methane to LH2 would be smaller advantage.As you said, trading density to ISP.

Hi guys, after reading your blogs about the use of different blow down gases..eg helium as oppossed to nitrogen can someone out there please further comment on this subject. We are building a LSR car using a LR-105 lox/kero injector and would like to use nitrogen as our blow down gas. Our total pressurization time is around 25 seconds for a run with a max 30 second pre pressurization time ( total 55 seconds propellant pressure time) From reading your comments i considered the main difference between using He or N20 was the time that the gas was putting the lox under pressure, am i correct ? Would N20 do the same job on a short duration blow down ? We are going to operate this vehicle in a very remote location where He is hard to source and is very expensive.Thank youRosco McGlashan Western Australia www.aussieinvader.com

On a SpaceX thread the possibility of using gaseous hydrogen dissolved in liquid methane came up. It looks as if this is impractical since the solubility is far too low. Would it be possible to use slush methane in liquid hydrogen and if so would this be more practical than slush hydrogen?

I'm struggling a bit to understand the advantage of doing this. Is the hope that dissolved methane would add little to the volume, thereby increasing volume-specific impulse (at the cost of weight-specific impulse)?

Coming from Elon's Reddit AMA video responses:"We can certainly improve on the chemical propulsion that has been done thus far, and I think probably a very high efficiency light hydrocarbon that uses predominantly methane is probably a good way to go, and I think that's something SpaceX will end up working on."

I was wondering what Elon was meaning and my idea was:Obtaining a (still unknown) mixture of liquid hydrocarbons that enables storing hydrogen at mild cryo temperature and low pressure would be a great advantage.Obviously adding solid methane to LH2 would be smaller advantage.As you said, trading density to ISP.

It would be fantastic if hydrogen were significantly soluble in some combination of light hydrocarbons, but I think that's unlikely, given the low solubility of hydrogen in methane. I would guess that Musk's remark refers to an engine running on a mixture of methane and either propane or proplylene or possibly a more exotic hydrocarbon.

Re dissolving or suspending solid methane in hydrogen, if feasible it certainly permits a higher volume-specific impulse, at a cost. I remain skeptical that it's worth trouble, when there are easier ways to boost volume-specific impulse.

I wasn't sure where to put this, and thought it might be useful to someone else.

This is a spreadsheet I built to answer a couple of questions of mine. It allows you to determine payload mass fraction for a given rocket based on a set of general performance metrics. You do not need to know any masses beforehand, but can calculate them afterward based on your desired payload mass.

It is currently set up to show why NTRs are not terribly useful for ground launch, which is the question I was curious about.

Most of the parameters are pretty self explanatory, but a couple notes:

Vehicle T/W and Engine T/W need to be consistent with each other. You can use lunar, Earth, Mars, Europa, or whatever the heck you want, because the gravity constant will cancel out. Just make sure you are using the same gravity constant for both T/W ratios.

Aux Structure Mass Fraction is a fudge factor to assign a percentage of vehicle mass for things other than engines, tanks, payload, or propellant.

Tank Density is the mass of a tank, per unit volume enclosed, in kg/m^3. As a few points of reference, the Shuttle SLWT is about 12 kg/m^3. A modern Centaur is about 25 kg/m^3. The Shuttle SRBs (as much as they were tanks...) were about 160 kg/m^3.

So the first thing you do when you're launching is loft yourself above the atmosphere, so it doesn't rob your v and crash you back down to Earth. Then you build up a bunch of v. (These are obviously happening concurrently e.g. with less-lofted launches.) I've seen the awesome Earth -> LEO chart circulating around here recently, and it lists that as costing 9.5 m/s. Does that include lofting costs? If so, how much of it? If not, how much extra dv-equivalent does it cost to get out of atmo? If I'm not asking the question quite right from my armchair, what is the right question (and answer)?

So the first thing you do when you're launching is loft yourself above the atmosphere, so it doesn't rob your v and crash you back down to Earth. Then you build up a bunch of v. (These are obviously happening concurrently e.g. with less-lofted launches.) I've seen the awesome Earth -> LEO chart circulating around here recently, and it lists that as costing 9.5 m/s. Does that include lofting costs?

Oh true, gravity hits you while you're trying to get up to altitude. Thx for the nomenclature. I meant to ask about losses to atmo friction + gravity loss, is that meaningfully more than gravity loss alone? How much are the two together? (approximately is cool. I'm not launching rockets here.)

Edit: or does gravity loss refer to atmo loss too? Since I guess you wouldn't need it if you were sans atmo.

What's the advantage of having multiple combustion chambers and nozzles on a given turbo machinery assembly instead of one larger chamber and nozzle? I know of one possible advantage on regenerative nozzles, but I'm thinking of other types such as on the Soyuz.

What's the advantage of having multiple combustion chambers and nozzles on a given turbo machinery assembly instead of one larger chamber and nozzle? I know of one possible advantage on regenerative nozzles, but I'm thinking of other types such as on the Soyuz.

For what I know, the bigger nozzles have bigger instabilities problems. Thus, it's easier to make more smaller ones, than a single bigger one. A second issue is that it's also shorter (smaller interstage). And a third issue is that since total thrust is just pressure times throat surface, but pressure vessels require proportionally more material for bigger diameters, you have thrust scaling quadratically, and weight cubically. Of course I suspect that the Russians do it because they design for low cost manufacturing. I.e. they have to tolerate very sloppy quality control. For a space application, obviously.